Gas turbine engines are generally used in a wide range of applications, such as aircraft engines and auxiliary power units. In a gas turbine engine, air is compressed in a compressor, and mixed with fuel and ignited in a combustor to generate hot combustion gases, which flow downstream into a turbine section. In a typical configuration, the turbine section includes rows of airfoils, such as stator vanes and rotor blades, disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled in turn to a main engine shaft. Hot combustion gases are delivered from the engine combustor to the annular hot gas flow path, thus resulting in rotary driving of the rotor disks to provide an engine output.
Due to the high temperatures in many gas turbine engine applications, it is desirable to regulate the operating temperature of certain engine components, particularly those within the mainstream hot gas flow path in order to prevent overheating and potential mechanical issues attributable thereto. As such, it is desirable to cool the rotor blades and stator vanes to prevent or reduce adverse impact and extend useful life. Mechanisms for cooling turbine rotor blades include ducting cooling air through internal passages and then venting the cooling air through holes formed in the airfoil. Internal and film cooling techniques attempt to maintain temperatures that are suitable for material and stress level. However, given the high temperature of engine operation, cooling remains a challenge, particularly in areas such as the turbine blade tips.
Accordingly, it is desirable to provide gas turbine engines with improved blade tip cooling. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.